Multi-Box Wing Spar And Skin

ABSTRACT

Apparatus and methods provide for the construction of wing sections having multi-box wing spars. According to one aspect of the disclosure provided herein, an aircraft wing may be constructed by applying one or more layers of composite material onto a plurality of wing mandrels. The wing mandrels, when placed together, form the shape of the wing. After the layers of composite material are applied to each individual mandrel, the mandrels are abutted together. Thereafter, the mandrels are compressed using wing surface tooling applied to the plurality of wing mandrels. The composite material is thereafter cured. After curing, the wing surface tooling and mandrels are removed, resulting in a wing having multi-box wing spars and skin.

BACKGROUND

Specific conventional aircraft wing construction varies frommanufacturer to manufacturer, but typically has several manufacturingprocesses in common. One or more wing spars that run the entire lengthof the wing are placed and secured in a wing construction jig. One ormore ribs are attached to the wing spars to give additional support tothe wing. After attaching the ribs to the wing spars, a series of wingstringers are coupled to the wing spars, giving additional structuralsupport as well as providing support to a wing skin. After theinstallation of one or more other features, such as fuel tanks,electronics, etc., as needed, the aircraft wing skin stiffened withstringers is attached to the wing spars and ribs. The wing skin may beattached using various methods, including, but not limited to, the useof rivets or other fasteners. Thereafter, other aircraft wing componentsare attached to the wing assembly, such as wing flaps, ailerons attachedto an aft spar, as well as forward and aft wing control surfaces.

Conventional techniques for constructing wings may use a relativelysignificant number of parts and may be a time-consuming, laboriousprocess. The number of parts may increase the weight of the wing as wellas the complexity of building the wing.

It is with respect to these considerations and others that thedisclosure made herein is presented.

SUMMARY

It should be appreciated that this Summary is provided to introduce aselection of concepts in a simplified form that are further describedbelow in the Detailed Description. This Summary is not intended to beused to limit the scope of the claimed subject matter.

Apparatus and methods provide for multi-box wing spars and skin usingone or more forming mandrels. According to one aspect of the disclosureprovided herein, an aircraft wing may be constructed by applying one ormore layers of composite material onto a plurality of wing mandrels. Thewing mandrels, when placed together, form the shape of the wing. Afterthe layers of composite material are applied to each individual mandrel,the mandrels are abutted together. Additional material may be added toform all or part of the skin of the wing or the top and/or bottom of themulti-box wing spars. Thereafter, the mandrels are compressed using wingsurface tooling applied to the plurality of wing mandrels. In someexamples, during mandrel compression, composite material may betensioned to straighten fibers in the composite material. The compositematerial is thereafter cured. After curing, the wing surface tooling andmandrels are removed, resulting in a wing having multi-box wing sparsand skins. If needed, one or more ribs are installed within themulti-box wing spar to provide for additional support.

According to another aspect, a wing may comprise several multi-box wingspars for attaching the wing to the fuselage of an aircraft. The wingmay have an upper surface, lower surface, and several wing segments. Themulti-box wing spars, upper wing surface and lower wing surface may beco-cured or co-bonded composite layers formed from substantiallycontinuous fibers.

According to a still further aspect, a system for forming an aircraftwing may include several mandrels shaped according to the upper, lower,forward and aft surfaces of the aircraft wing. The system may alsoinclude a compression apparatus for compressing the several mandrelstogether to cure layers of composite material on the mandrels. Thesystem may also include a tension block for maintaining tension on thecomposite material.

The features, functions, and advantages that have been discussed can beachieved independently in various configurations of the presentdisclosure or may be combined in yet other configurations, furtherdetails of which can be seen with reference to the following descriptionand drawings.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a top perspective view of an exemplary mandrel that may beused to form multi-box wing spars, according to various configurationspresented herein;

FIG. 2 is a top perspective view of an exemplary mandrel that may beused to form a wing illustrating the application of a filament woundbias ply to the mandrel, according to various configurations presentedherein;

FIG. 3 is a top perspective view of an exemplary mandrel that may beused to form a wing illustrating the application of a uni-ply to themandrel, according to various configurations presented herein;

FIG. 4 is a top perspective view of an exemplary mandrel that may beused to form a wing illustrating a finished layup on the mandrel,according to various configurations presented herein;

FIG. 5 is a top perspective view of a plurality of exemplary mandrelsthat may be used to form a wing, according to various configurationspresented herein;

FIG. 6 is a top perspective view of a plurality of exemplary mandrelsthat are abutted to each other, according to various configurationspresented herein;

FIG. 7 is a top perspective view of wing surface tooling prior tocompression, according to various configurations presented herein;

FIG. 8 is a top perspective view of wing surface tooling duringcompression, according to various configurations presented herein;

FIG. 9 is a top perspective view of wing surface tooling being removedafter curing, according to various configurations presented herein;

FIG. 10 is a top perspective view of a wing having multi-box wing sparsand skin after the removal of the wing surface tooling and mandrels,according to various configurations presented herein;

FIG. 11 is a top perspective view of a wing having multi-box wing sparsand skin after the removal of the wing surface tooling and mandrels,further illustrating the installation of a vertical rib, according tovarious configurations presented herein;

FIG. 12 is a top perspective view of a fuselage section with two wingshaving multi-box wing spars, according to various configurationspresented herein; and

FIG. 13 is an illustrative routine for manufacturing multi-box wingspars, according to various configurations presented herein.

DETAILED DESCRIPTION

The following detailed description provides for wings having multi-boxwing spars. As discussed briefly above, conventional aircraft wings aretypically constructed using multiple components, including one or morespars, ribs and wing stringers, each performing individual functions.After the wing is constructed, the wing is typically attached to a wingbox on the fuselage of the aircraft. In conventional aircraft, the wingbox is a reinforced, structural component on the aircraft's fuselage towhich the wings are attached. Conventional aircraft wing constructioncan be a time-consuming and costly process. Further, if the wing isconstructed using certain materials, including polymers or, especially,composite materials, the multiple sections of the wing reduce theavailability of relatively long lengths of un-broken material, reducingthe strength of the polymers and/or composite materials. For example,when using carbon fiber reinforced thermoplastics, if the surfacecontains multiple breaks in the fibers, the benefit of using thecomposite can be reduced, as additional reinforcing structures ormaterial may need to be used to make up for the loss in structuralrigidity.

Utilizing the concepts described herein, an aircraft wing may beconstructed using a series of forming mandrels that when placed togetherform the shape of the wing. Composite materials, or other suitablematerials, may be applied to each forming mandrel, and thereaftercompressed and cured to form a wing having multi-box wing spars. Inother configurations, one or more layers of composite materials may befully or partially cured prior to application of the layers to one ormore parts of the forming mandrel. In that configuration, one or morepartially or fully cured layers of composite material may be co-bondedwith other partially or fully cured layers of composite materials. Itshould further be appreciated that the concepts described hereinrelating to an aircraft wing may also be used for other aircraftcomponents, such as a vertical or horizontal stabilizer, withoutdeparting from the scope of this disclosure and the accompanying claims.

In the following detailed description, references are made to theaccompanying drawings that form a part hereof, and which are shown byway of illustration, specific configurations, or examples. Referring nowto the drawings, in which like numerals represent like elements throughthe several figures, the manufacture of wings having multi-box wingspars will be described. It should be appreciated that a multi-box wingspar constructed according to various configurations disclosed hereinmay have one or more spars and one or more skins, the present disclosureof which is not limited to any number of spars or skins.

Turning to FIG. 1, is a top perspective view of an exemplary mandrelthat may be used to form a wing is illustrated. Forming (or wing)mandrel 100 may have upper surface layer 102 that is shaped according toan upper surface layer of an aircraft wing (shown by way of example inFIG. 12). Forming (or wing) mandrel 100 may also have lower surfacelayer 104 that is shaped according to a lower surface layer of anaircraft wing. Upper surface layer 102 and lower surface layer 104 maybe shaped so that when one or more layers of composite material, such ascarbon fiber reinforced thermoplastic, is applied to mandrel 100, theresulting shape is in the shape of an aircraft wing. Further, uppersurface layer 102 and lower surface layer 104 may be shaped so thatthere are little to no bends or breaks in the fibers of the compositematerial, providing for increased rigidity from the composite material.

As will be described in more detail below, when forming a structureusing composite materials, it is typically preferable to not only avoidunnecessary bends or breaks in the material, but also to maintain thestraightness of the fibers running in one or more portions of thematerial. In typical composite materials, it may be preferable to havestraight fibers in the composite material matrix. It should beappreciated that the present disclosure is not limited to the use ofstraight fibers in a composite matrix. Various configurations of thepresent disclosure may be implemented in the construction of wingshaving composite matrices using curved or straight fibers, orcombinations thereof.

If it is desired or necessary to straighten the fibers in a compositefiber matrix prior to curing, mandrel 100 may also have tension blocks106 and 108. Tension blocks 106 and 108 may be used separately or inconjunction with one another to “pull” the fibers of a composite matrix,thus providing for a straightened fiber matrix. In some configurations,fibers in a composite matrix (described in more detail below) may beattached to one or both of tension blocks 106 and 108 or may be formedaround tension blocks 106 and/or 108. Tension blocks 106 and 108 may beconfigured to extend outwards from mandrel 100 at various pressures toprovide for a tension on the fibers in the fiber matrix.

A configuration of the construction of the fiber matrix, as discussedbriefly above, and the forming of a wing having multi-box wing spars arenow described in relation to FIGS. 2-12. In FIG. 2, a first layer offilament wound bias ply 200 (illustrated in a cross-hatch pattern overthe surface of mandrel 100) is wrapped or applied to mandrel 100. Ifneeded or desired, filament wound bias ply 200 may have tension appliedthereto using tension block 106 and/or tension block 108 to helpstraighten the fibers in filament wound bias ply 200.

After the filament wound bias ply 200 is applied to mandrel 100, auni-ply layer is applied, as shown in FIG. 3. Uni-ply layer 300 isapplied to mandrel 100 and may be tightened using tension block 106and/or tension block 108. It should be understood that the presentdisclosure is not limited to any specific configuration of bias-ply oruni-ply. For example, one or more layers of a bias-ply layer may beadded prior to the addition of a uni-ply layer. In the same manner, oneor more layers of uni-ply may be added between the applications of theone or more bias-ply layers.

Further it should be understood that the present disclosure is notlimited to any number of layers of either type of layer, as variouscombinations may be used to achieve structural or cost goals. Forexample, and not by way of limitation, it may be desirable or necessaryto add sufficient layers of wound bias-ply and/or uni-ply to achieve adesired wing thickness or structural rigidity. Additionally, it shouldbe understood that the present disclosure is not limited to a layerhaving a single type of ply, as some configurations may use acombination of bias- and uni-ply within the same layer. Variouscombinations may be used according to various configurations withoutdeparting from the scope of this disclosure and the accompanying claims.

FIG. 4 illustrates a completed composite matrix 400 on mandrel 100.Composite matrix 400 may be formed using various layering andapplication techniques, such as, by way of example, the method describedabove in relation to FIGS. 2 and 3.

FIG. 5 is a top perspective view of a series of mandrels having fullyformed composite matrices applied thereon. Mandrel apparatus 500 hasindividual mandrels 500 a-d. Mandrels 500 a-d have disposed thereoncomposite matrix 502, having individual composite matrices 502 a-d.Composite matrix 502 may be formed from one or more layers, variouscombinations of plies, and may be fully or partially uncured at thispoint. As illustrated in FIG. 5, the general shape of an aircraft wingmay be seen when viewing mandrel apparatus 500.

According to various configurations, after composite matrix 502 isapplied to mandrel apparatus 500, mandrels 500 a-d are abutted againsteach other, as shown in FIG. 6. Individual mandrels (illustrated by wayof example as mandrels 500 a-d in FIG. 5) are abutted to form acontiguous, multi-box wing spar layup, having composite matrix 502,which is formed from multiple composite matrices (illustrated by way ofexample as composite matrices 502 a-d in FIG. 5).

In order to partially or fully cure and form the multi-box wing sparconfiguration according to various configurations of the presentdisclosure, a curing system may be used. As discussed above, one or morelayers of composite material may be fully or partially cured prior touse in a forming mandrel. In that configuration, the partially or fullycured composite layers may be co-bonded to other partially or fullycured composite layers using one or more layers of adhesive to securethe layers of composite material within a composite matrix. An exampleof a system for forming multi-box wing spars is shown in FIG. 7. Afterthe individual mandrels forming mandrel apparatus 500 are abutted toeach other, thus forming composite matrix 502 from a series ofindividual composite matrices, a series of surface tools may be appliedto the various surfaces of mandrel apparatus 500. It should beappreciated that additional material may be added across the surface ofthe spar after the mandrels that form mandrel apparatus 500 are abutted.The additional material may be used to form the skin of the wing,reinforce the composite material already in place, or provide forvarious aerodynamic or physical properties, by way of example.

Various processes for adding the additional bias-ply and uni-ply areknown to those in the art, to which the various configurations disclosedherein are not dependent on any one particular method of applyingcomposite materials. In one configuration, a curable upper wing skin, acurable lower wing skin, a curable leading wing edge and a curabletrailing wing edge may be applied after the initial plies are added tothe composite matrix 502, prior to the application of surface tools tothe composite matrix 502.

Compression apparatus 504 has forward skin surface tool 506, lower skinsurface tool 508, aft skin surface tool 510 and top skin surface tool512. It should be further appreciated that not all of the material addedis “curable” material, as non-curable material may be added to compositematrix 502. Surface tools 506, 508, 510 and 512 are individually orcollectively compressed, thus applying pressure, onto the respectivesurfaces of mandrel apparatus 500 to help form and cure composite matrix502. In some configurations, heating element 514 may be applied to oneor more of surface tools 506-512. The combination of pressure and heatmay fully or partially cure composite matrix 502 in a desired amount oftime or may provide for additional structural rigidity. Heating element514 may use various means of applying heat to composite matrix 502,including steam and electrical current. FIG. 8 illustrates compressionapparatus 504 in a compressed state, with mandrel apparatus 500 shownoutside of compression apparatus 504.

Once composite matrix 502 is cured to a desired level, compressionapparatus 504 is removed, illustrated in greater detail in FIG. 9.Compression apparatus 504 surface tools 506, 508, 510 and 512 areremoved from the surface of now-cured composite matrix 502 andindividual mandrels of mandrel apparatus 500 are extracted fromcomposite matrix 502. The resulting structure is illustrated in FIG. 10.It should be appreciated that the present disclosure is not limited tofully curing composite matrix 502, as it may be desired or necessary insome configurations to remove the curing mechanisms (e.g. compressionapparatus 504 or heating element 514) prior to composite matrix 502being fully cured. Various degrees of curing may be used according tovarious configurations of the present disclosure without departing fromthe scope of this disclosure and the accompanying claims.

FIG. 10 is a top perspective illustration showing composite matrix 502with multi-box wing spar. By using composite materials formed overabutting mandrels, composite matrix 502 has disposed therein spars 600a-e that extend internally to composite matrix 502 along axis X-Y, thusforming multi-box wing spars. By using a mandrel apparatus, such asmandrel apparatus 500 of FIG. 7, it can be seen that composite matrix502 may be formed having fibers disposed therein that can bestraightened and uncut (or undesirably terminated). After curing,composite matrix 502 may be considered a singular, contiguous structure.If it is desirable or necessary to further reinforce composite matrix502, one or more ribs may be installed in composite matrix 502, anexample of which is shown by rib 700 in FIG. 11.

Further, utilizing the concepts described herein, a wing formedaccording to various configurations disclosed herein may be coupled to afuselage section of an aircraft without the need for a conventional wingbox. Exemplary techniques are described in copending patent applicationentitled, “Vertically Integrated Stringers,” filed on Nov. 26, 2012,which is hereby incorporated herein in its entirety. FIG. 12 illustratesone such configuration in which a wing formed according to thetechniques described herein is attached to a fuselage without the use ofa traditional wing box. It should be appreciated that the conceptspresented herein may also be used to form a wing according to thetechniques described herein to be attached to a traditional wing box.

Composite matrices 800 and 802, which are constructed according tovarious configurations disclosed herein, have multiple wing spars thatcan be coupled to fuselage section 804. It should be appreciated thatcomposite matrices 800 and 802 may be formed in various shapes withvarious features, the present disclosure of which is not limited to anyone particular configuration. Exemplary wing spar 806 is identified inFIG. 12 for the sake of clarity, though it should be understood thatcomposite matrices 800 and 802 may have additional wing spars. Exemplarywing spar 806 may have elliptical aperture 808.

Depending on the angular displacement between composite matrices 800/802and fuselage section 804, elliptical aperture 808 may vary incircumference and shape, i.e. the foci of elliptical aperture 808 maychange as well as the radii. For example, in a straight-wing profileaircraft in which spar 806 may be affixed to fuselage section 804 atapproximately a 90 degree angle, elliptical aperture 808 may becircular. In another example, such as the one illustrated in FIG. 12,spar 806 may be attached to fuselage section 804 in a swept-wingprofile. Thus, elliptical aperture 808 may be more oval in shape inorder to provide for interior space in the aircraft and to be attachedto the fuselage circumferentially. One or more circumferential fuselagestringers, such as stringers disclosed in copending application entitled“Vertically Integrated Stringers” and identified as circumferentialstringers 810, may provide additional structural support to fuselagesection 804. The multi-box wing spars formed by composite matrices 800and/or 802 may be attached to one or more beams of an aircraft fuselage,such as crown beam section 812. It should be appreciated that fuselagesmay have one or more types of beams including, but not limited to, crownbeam 812 or a keel beam (not shown).

FIG. 12 also illustrates the various sections of a wing that may beformed using various configurations disclosed herein. Composite matrix802 is illustrated as having a leading wing edge 814, trailing wing edge816, upper surface layer 818 and lower surface layer 820. One or more ofthe leading wing edge 814, the trailing wing edge 816, the upper surfacelayer 818 and the lower surface layer 820 may be curable or bondableaccording to various configurations disclosed herein. Further, one ormore of the leading wing edge 814, the trailing wing edge 816, the uppersurface layer 818 and the lower surface layer 820 may be formedseparately from the others and attached afterwards. In someconfigurations, the leading wing edge 814 and/or the trailing wing edge816 may be formed with the upper surface layer 818 and/or the lowersurface layer 820. Thus, in a compression apparatus, such as compressionapparatus 504 of FIG. 7, the leading wing edge 814 may be a forward skinformed using the forward skin surface tool 506 of FIG. 7 and thetrailing wing edge 816 may be an aft skin surface formed using the aftskin surface tool 510.

Turning now to FIG. 13, an illustrative routine 900 for constructingmulti-box wing spars is described in detail. Unless otherwise indicated,it should be appreciated that more or fewer operations may be performedthan shown in the figures and described herein. Additionally, unlessotherwise indicated, these operations may also be performed in adifferent order than those described herein.

Routine 900 begins at operation 902, where one or more bias- and/oruni-plies are applied to a series of forming mandrels. In someconfigurations, the fibers in the bias- and/or uni-plies can betightened through the use of one or more tension blocks on the mandrel.From operation 902, routine 900 continues to operation 904, whereby theforming mandrels are abutted to each other to create a multi-box wingspar layup. In some configurations, the multi-box wing spar layupcomprises a composite matrix formed from one or more layers of the bias-and/or uni-plies. As noted above, additional plies may be added atvarious stages of the forming process.

From operation 904, routine 900 continues to decision 906, wherein adetermination is made if additional layers of uni-ply or bias-ply are tobe added prior curing of the composite matrix. In one configuration, itmay be desirable to form and cure together the multi-box wing spars andone or more portions of the wing skin. In another configuration, adesired wing thickness or structural rigidity may require thatadditional plies be added. If the determination 906 is that additionalply layers are to be added, routine 900 continues to operation 908,wherein the additional layers are applied to the layup.

If it was determined 906 that no additional layers to the layup are tobe applied 908, or after the additional layers to the layup have beenapplied 908, routine 900 continues to operation 910, whereby surfacetooling is applied (abutted) to the various surfaces of the compositematrix. The surface tooling, in some configurations, may serve severalfunctions. For example, surface tooling may have one or more surfacesconfigured to create certain shapes in the surface of the compositematrix. Surface tooling may also be used to apply pressure and/or heatto a composite matrix to cure the composite matrix as well as, in someexamples, provide for debulking of the composite matrix during layup.

From operation 910, routine 900 continues to operation 912, whereby thesurface tooling is compressed onto the composite matrix to being thecuring process. In some configurations, it may be desirable to, inaddition to pressure, apply heat to one or more surface tools, heatingvarious surfaces of the composite matrix. Thus, operation 912 may alsoinclude a heating operation.

From operation 912, routine 900 continues to operation 914, whereby thecomposite matrix in the multi-box wing spar layup is cured. In somefurther configurations, it may be desirable at operation 912 and/oroperation 914 to apply tension to the plies within the composite matrixfrom one or both ends of the composite matrix to reduce the amount ofwrinkles of fibers within the composite matrix and to increase thestraightness of the fibers within the composite matrix. Once the curingcycle is completed, routine 900 continues to operation 916, whereby thesurface tooling (and heat) is removed from the composite matrix.Further, the mandrels are extracted from the composite matrix, formingwing sections having multi-box wing spars.

Based on the foregoing, it should be appreciated that technologies forconstructing wing sections having multi-box wing spars have beenpresented herein. The subject matter described above is provided by wayof illustration only and should not be construed as limiting. Variousmodifications and changes may be made to the subject matter describedherein without following the example configurations and applicationsillustrated and described, and without departing from the true spiritand scope of the present disclosure, which is set forth in the followingclaims.

What is claimed is:
 1. A method of manufacturing an aircraft wing,comprising: applying a plurality of filament wound bias-ply and uni-plyon each of a plurality of separated wing mandrels; abutting theplurality of separated wing mandrels to create a multi-box wing sparlayup; abutting a plurality of wing surface tooling to the multi-boxwing spar layup; compressing the plurality of wing surface tooling toapply pressure to the multi-box wing spar layup; and curing themulti-box wing spar layup to form the aircraft wing having a pluralityof multi-box wing spars.
 2. The method of claim 1, wherein curing themulti-box wing spar layup further comprises applying heat to themulti-box wing spar layup.
 3. The method of claim 1, further comprisingremoving each of the plurality of separated wing mandrels after curingthe multi-box wing spar layup.
 4. The method of claim 1, whereinapplying a plurality of filament wound bias-ply and uni-ply comprisesadding sufficient layers of wound bias-ply or uni-ply to achieve adesired wing thickness or structural rigidity.
 5. The method of claim 1,further comprising laying-up a curable upper wing skin and a curablelower wing skin after applying a plurality of filament wound bias-plyand uni-ply, wherein curing the multi-box wing spar layup co-cures themulti-box wing spar layup, the upper wing skin and the lower wing skin.6. The method of claim 1, further comprising installing a plurality ofrib segments after curing the multi-box wing spar layup.
 7. The methodof claim 1, further comprising installing a leading wing edge or atrailing wing edge after curing the multi-box wing spar layup.
 8. Themethod of claim 1, further comprising laying-up a curable leading wingedge and a curable trailing wing edge after applying the plurality offilament wound bias-ply and uni-ply, wherein curing the multi-box wingspar layup co-cures the multi-box wing spar layup, the leading wing edgeand the trailing wing edge.
 9. The method of claim 1, whereincompressing the plurality of wing surface tooling further comprisesapplying tension to the plurality of filament wound bias-ply and uni-plyto minimize wrinkling of a plurality of fibers therein.
 10. A wing,comprising: a plurality of multi-box wing spars attached to at least onebeam of an aircraft fuselage; an upper wing surface; a lower wingsurface; and a plurality of wing segments, wherein the plurality ofmulti-box wing spars, upper wing surface, and lower wing surface arecomposite layers comprising substantially continuous fibers.
 11. Thewing of claim 10, wherein the wing further comprises a plurality of ribsdisposed within the plurality of multi-box wing spars.
 12. The wing ofclaim 10, wherein the wing further comprises a leading wing edge and atrailing wing edge comprising composite layers co-cured with theplurality of multi-box wing spars, upper wing surface, and lower wingsurface.
 13. The wing of claim 10, wherein at least one of the multi-boxwing spars comprises an elliptical aperture.
 14. The wing of claim 13,wherein an outer surface of the elliptical aperture is proximate to aninside surface of the aircraft fuselage.
 15. The wing of claim 13,wherein foci of the elliptical aperture provide an angular displacementbetween the aircraft fuselage and the wing, wherein the angulardisplacement provides for a straight-wing or swept-wing profile.
 16. Thewing of claim 10, wherein at least one of the plurality of multi-boxwing spars is attached to at least one of a plurality of multi-box wingspars of a second wing.
 17. The wing of claim 10, further comprising awing skin co-cured with the plurality of multi-box wing spars.
 18. Thewing of claim 10, wherein the composite layers comprising substantiallycontinuous fibers are co-cured or co-bonded composite layers.
 19. Asystem for forming an aircraft wing, comprising: a plurality of mandrelscomprising: an upper surface layer shaped according to an upper surfacelayer of the aircraft wing; a lower surface layer shaped according to alower surface layer of the aircraft wing; and a forward surface layerand an aft surface layer abutted with one or more of the plurality ofmandrels, wherein the upper surface layers of the plurality of mandrelsare shaped according to an upper surface of the aircraft wing and thelower surface layers of the plurality of mandrels are shaped accordingto a lower surface of the aircraft wing; a tension block for maintainingtension on one or more layers of composite material forming the aircraftwing; and a compression apparatus for mechanically curing the one ormore layers of composite material forming the aircraft wing, thecompression apparatus comprising: an aft skin surface tool for applyingpressure to the aft surface layer of the plurality of mandrels; a topskin surface tool for applying pressure to the upper surface layer ofthe plurality of mandrels; a lower skin surface tool for applyingpressure to the lower surface layer of the plurality of mandrels; and aforward skin surface tool for applying pressure to the forward surfacelayer of the plurality of mandrels.
 20. The system of claim 19, furthercomprising a heating element for thermally curing the one or more layersof composite material forming the aircraft wing.